The use and configuration of bleed assemblies are well known in gas turbine engines and are usually used to improve engine operability, particularly for the engine's compressors. In use, heated air at high pressure passes from a compressor, through a bleed valve and via a diffuser screen into a main gas stream. The compressor may be, for example, an intermediate or high pressure compressor and the bled gas stream may have a temperature of up to 400° C. The diffuser screen is usually a domed plate comprising a pattern of through-holes to enhance mixing with the main gas stream, which is typically a cooler bypass flow.
The total number, size and spacing of the through-holes of the diffuser screen are governed by performance and acoustic considerations. For example, diffuser screens can be configured to (a) attenuate noise produced within the bleed valve, (b) produce small separate jets of bleed air (rather than one large one) to increase the jets' noise frequency, which is then better attenuated by acoustic liners within the bypass duct and atmosphere, and (c) improve mixing of the hot gas flowing out of the bleed assembly with the cooler bypass flow in order to limit/prevent thermal damage to the engine nacelle and other components.
FIG. 1 shows a schematic longitudinal cross-section through a ducted fan gas turbine engine generally indicated at 10 and having a principal and rotational axis 11. The engine comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, and intermediate pressure turbine 18, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines the intake 12, a bypass duct 22 and an exhaust nozzle 23.
The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two airflows: a first airflow A into the intermediate pressure compressor 14 and a second airflow B which passes through the bypass duct 22 to provide propulsive thrust.
During engine operations, and particularly when changing rotational speed at low power, it is important to ensure that the pressure ratio across each compressor 14, 15 remains below a critical working point, otherwise the engine 10 can surge and flow through the engine breaks down. This can cause damage to the engine's components as well as aircraft handling problems.
To maintain a preferred pressure difference across both or just one of the compressors 14, 15, bleed assemblies 30 are provided to release pressure from an upstream part of a compressor.
FIG. 2 is a schematic cross-section showing in more detail bleed assemblies 30 associated with the intermediate pressure compressor 14 and high pressure compressor 15. Each bleed assembly comprises an inlet 31 and a bleed valve 32, a duct 34 and a diffuser screen 36. Airflows C and D, which are parts of core engine airflow A, may be diverted through the respective bleed assemblies, such that each airflow C, D enters the inlet, passes through the bleed valve and is channelled by the duct to the diffuser screen. Airflows C and D are then exhausted into the bypass duct 22 where they mix with bypass airflow B as hereinbefore described. There is usually an annular array of bleed valves around the core engine's casing 27.
However, a problem can arise that diffuser screens may rupture.